Operation of turbine engines is well known. Referring to FIG. 1, a gas turbine engine is generally indicated at 10 and comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high pressure compressor 14, a combustor 15, a turbine arrangement comprising a high pressure turbine 16, an intermediate pressure turbine 17 and a low pressure turbine 18, and an exhaust nozzle 19.
The gas turbine engine 10 operates in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 which produces two air flows: a first air flow into the intermediate pressure compressor 13 and a second air flow which provides propulsive thrust. The intermediate pressure compressor compresses the air flow directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.
The compressed air exhausted from the high pressure compressor 14 is directed into the combustor 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive, the high, intermediate and low pressure turbines 16, 17 and 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low pressure turbines 16, 17 and 18 respectively drive the high and intermediate pressure compressors 14 and 13 and the fan 12 by suitable interconnecting shafts.
In view of the above it will be appreciated that there is leakage about the peripheral tips of rotor blades of the turbines 16, 17 and 18 during operation and this leakage reduces efficiency with regard to the engine as well as causing heating problems due to the hot combustion gas nature of the leakage flow about the tip. It is known to inhibit leakage using shrouds on either side of the rotor tip, but generally such shrouding adds significantly to weight and therefore is impractical within most turbine engines particularly if utilised in aircraft. Unshrouded rotor tip leakage prevention is also known, and an example is outlined in U.S. Pat. No. 6,142,739 (Rolls Royce plc). Essentially, in this unshrouded rotor tip, a gutter is defined at the tip within which leakage flow is captured. Such leakage capture is normally through induced vortex effects. Nevertheless, as indicated previously, such captured leakage will generally be of a hot combustion gas such that there may be significant heating problems with respect to turbine tips which in themselves are generally formed from thinner sections of material.
Modern gas turbines operate with high turbine entry temperatures to achieve high thermal efficiencies. These temperatures are limited by the turbine vane and blade materials. Cooling of these components is now needed to allow their operating temperatures to exceed the material's melting point without affecting the vane and blade integrity.
A large number of cooling systems are now applied to modern high temperature gas turbine vanes and blades. Cooling is achieved using relatively cool air bled from the upstream compressor system and is arranged to bypass the combustion chamber between the last compressor stage and first turbine stage. This air is introduced into the turbine vanes and blades where cooling is effected by a combination of internal convective cooling and external film cooling.
In film cooling a protective blanket of cooling air is ejected onto the external surface of the turbine vane or blade, from internal passages within the aerofoils, by means of holes or slots in the surface. The aim is to minimise the external heat transfer from the hot gas stream into the component surface.
In convective cooling the air is passed through passages within the aerofoil which cools the metal since the air temperature is below that of the metal. Effectively the turbine component itself acts as a heat exchanger.
U.S. Pat. No. 6,142,739 (Rolls-Royce plc) illustrates one unshrouded turbine blade tip. In short, a gutter is provided along the top of that tip. FIG. 2 provides a simple isometric view of a rotor blade tip of a so called partial shroud or winglet type. Essentially, a rotor blade 20 has a blunt trailing edge 21 and the partial shroud 22 everywhere lies outside the envelope of the rotor blade 20 aerofoil beneath it. An internal or gutter channel 23 is provided which extends from a leading edge 24 to the trailing edge 21. This channel 23 is open at both ends and widens from a leading edge 24 to the trailing edge 21. The open ends to the channel 23 are provided in order to initiate cross-flow to enter to the gutter channel 23. It should be appreciated that the opening at the leading edge 24 to the gutter channel 23 provides an additional leakage path and so is kept as small as possible.
Of importance with regard to the present invention is that some of the leakage flow remains within the internal gutter channel 23 and does not reach the suction side S. Operationally this is important as it prevents mixing with the high velocity main turbine stream flow on the suction side S which would result in high aerodynamic losses. Unfortunately, lingering of leakage flows within the gutter channel 23 as indicated causes significant heating problems, etc.